Gas turbines and the modes of operation thereof are generally known. The air drawn in by a compressor of the gas turbine is compressed therein and afterwards mixed with fuel in a burner. Subsequently, the mixture flowing into a burning chamber burns to form a hot gas which subsequently flows through a turbine arranged downstream of the burning chamber and in the meantime, owing to the relaxation thereof, causes the rotor of the gas turbine to rotate. The rotation of the rotor drives, in addition to the compressor, also a generator which is linked to the rotor and converts the mechanical energy provided into electrical energy.
Both the compressor and the turbine are each composed of a plurality of successive stages each comprising two successive collars of blades. A turbine stage is composed of a guide blade collar formed by rotationally fixed guide blades and a moving blade collar arranged downstream thereof, whereas a compressor stage is composed of a moving blade collar and a guide blade collar arranged downstream thereof; viewed in each case in the flow direction of the medium flowing through. In the case of a single-shaft gas turbine, all moving blades are fixedly mounted on the common rotor.
The compressor stages, which are arranged in series, i.e. in axial succession, convey, owing to the moving blades revolving with the rotor, the drawn-in air from the input of the compressor in the direction of the compressor output, the air experiencing an incremental rise in pressure within each stage (or collar). The total rise in pressure over the compressor is the sum of all incremental pressure rises over each stage (or of all collars).
In a known manner, it can occur during operation of the gas turbine, in particular during operation of the compressor of the gas turbine, that, on approaching the stability limit, recirculation is increased as a result of defective flow and growing gap vortex. Within the compressor, this can cause a stall on one or more aerofoils, i.e. the flow of air in the main flow direction stops through a part of a compressor stage, as the energy transmitted from the rotor to the air is not sufficient to convey said air through the compressor stage and to establish the required pressure ratio of the compressor stage in question. The pressure ratio is the increase in pressure occurring over the relevant stage of the compressor, based on the input pressure of the respective stage. If the stall is not immediately counteracted, it can advance to form a rotating stall and possibly even lead to the entire flow of air through the compressor changing its direction; this is known as compressor pumping. This particularly critical operating state jeopardizes the blades and prevents a sufficient supply of compressor air to the burning chamber, so that a disturbed operation of the gas turbine must be diagnosed and the machine switched off immediately.
For this purpose, EP 0 719 907 A1, which seeks to counteract the described problem, discloses a structured boundary wall which lies opposite the tips of the moving blades. This structuring of the casing, known as the casing treatment, serves positively to influence the flow close to the gap for situations in which there is a risk of a stall on an aerofoil. Owing to the structuring, partial flows are bled from the main flow in the region of low flow velocities and subsequently returned to the main flow upstream of the bleed position. The air bled on the pressure side of the compressor blades in the tip region is supplied to the suction-side main flow of the compressor blade in question to prevent a stall which might occur there. The ducts guiding the partial flows are accordingly inclined relative to the machine axis or axis of rotation in such a way that—viewed in the direction of rotation of the rotor—the bleed position lies after the feed position at which the detached partial flow is returned to the main flow close to the gap. This is required so that, owing to the stagger angle and the tips, positioned obliquely relative to the direction of rotation, of the aerofoils, the partial flow can be guided beyond the aerofoil tip from the pressure side to the suction side. Thus, the longitudinal extension of the return flow duct is oriented substantially transversely to the straight line of the blade tip-side stagger angle, i.e. approximately parallel to the machine axis.
A similar device is known from EP 1 286 022 A1.
The aforementioned configurations have the drawback that the flow guidance of the partial flows is not optimal.
Furthermore, FR 2 325 830 discloses a compressor casing with grooves formed therein. These grooves are intended to prevent a stall of the limit flow and thus the pumping of the compressor, although the flow set by the grooves does not flow counter to the main flow, but rather with it.